1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine stator vane with near wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Turbine stator vanes typically use an impingement insert to direct impingement cooling air from a supply channel to the backside surface of a hot wall surface of the vane. Stator vanes can use inserts because they are non-rotating airfoils as opposed to rotor blades. FIG. 1 shows a prior art turbine vane with an insert that provides backside impingement cooling for the entire airfoil. The airfoil 11 includes a pressure side wall and a suction side wall extending between a leading edge region and a trailing edge region with a cooling supply cavity 13 formed between the walls. An insert tube 12 includes impingement holes 14 that direct impingement cooling air to selected sections of the airfoil walls to provide for the backside impingement cooling. A number of stand-offs 15 are positioned to secure the insert tube 12 in place within the cavity 13.
In operation, cooling air from the supply cavity 13 flows through the impingement holes 14 in parallel to produce impingement cooling for the backside surface of the airfoil walls. The spent impingement cooling air is then collected within a passage 16 formed between the insert tube 12 and the airfoil inner walls and channeled toward the trailing edge region where the cooling air is then discharged through a row of trailing edge exit holes 17 that can include pin fins to enhance the heat transfer from the trailing edge region metal to the cooling air.
The FIG. 1 prior art vane cooling circuit requires a relatively high cooling flow rate because of the parallel arrangement of impingement cooling holes. The cooling air is spread out very thin in order to cover the entire backside of the airfoil. With this arrangement, the cooling of the hot gas surface area is very low. In a low flow cooling design, the spacing in-between the impingement holes are so far apart that the areas between impingement holes are without backside impingement cooling. Also, a continuous impingement cooling channel will also produce a cross flow effect and therefore degrade the impingement heat transfer coefficient and reduce the overall cooling effectiveness. Plus, a relatively thick airfoil wall will increase the conduction path of the impingement cooling air that will reduce the thermal efficiency for the airfoil backside impingement cooling. Other embodiments can have a rib that extends across the cavity to form multiple cooling air cavities each with a separate impingement place or insert.